The desire for improved performance and fuel efficiency in turbofan and other turbine engines has produced continuous improvement in the specific thrust and specific fuel consumption of such engines. Specific thrust is determined by dividing the pounds of thrust produced by the turbine engine by the pounds of the total (core and bypass) mass flow of air the engine "swallows." Specific fuel consumption is determined by dividing the pounds of fuel consumed per hour by the turbine engine divided by the pounds of thrust produced by the engine.
Specific thrust may be increased and specific fuel consumption may be decreased by increasing the cycle pressure ratio ("CPR") of the turbine engine. CPR is determined by dividing the pressure of air exiting the high pressure compressor of the engine by the pressure of the air at the inlet of the turbine engine. For aircraft turbofan engines, the most advanced production engines for subsonic civilian and military aircraft have CPRs of about 40-45 and the most advanced production engines for supersonic military aircraft have CPRs of about 20-25.
Cycle analysis indicates increases in CPR beyond today's levels results in a further increase in turbine engine performance. Also, the full performance benefits of increased CPR may be realized when combined with increased turbine inlet temperature ("TIT"), i.e., the temperature of the gases at the inlet of the high pressure turbine downstream from the combustion chamber. Unfortunately, limitations in materials make it difficult to achieve such an increase in CPR and TIT.
There are two major locations at which the limitations in materials capability prevent further increases in CPR and TIT beyond the today's level. One of the locations is in the downstream half of the high pressure compressor. Although high temperature resistant materials such as nickel alloys are used today in this half for compressor blades, disks and conical, torque-transmitting diaphragm (or shaft), a further increase of core air temperature would reduce creep life of these components to an unacceptable extent.
The second major location where materials limitations prevent increases in CPR is the high pressure turbine. Today, the state-of-the-art level of TIT is 2,800.degree. F. to 3,200.degree. F. at maximum turbine engine load, e.g., during aircraft take-off. These values are achievable only by cooling the high pressure turbine components, such as nozzle vanes, blades, disks and shrouds which are made from high temperature resistant materials such as nickel alloys, with a portion of the high pressure compressor delivery air ("T3 air"). With this cooling, higher TIT can be accommodated. For example, U.S. Pat. No. 5,269,133 to Wallace describes a turbofan engine in which a relatively small portion (i.e., about 10%) of the T3 air is transported radially from the core passage through a heat exchanger in the bypass passage and then is returned radially to a manifold that delivers the relatively cooled air to the high pressure turbine.
However, increasing the compression ratio of the high pressure compressor in an attempt to increase CPR beyond today's level would result in a temperature increase in the high pressure compression delivery flow. This in turn would necessitate an increased delivery of cooling air flow to the high pressure turbine to maintain creep life of the turbine components at the today's level at current maximum TITs. An increase of turbine cooling air flow reduces the effective use of the core flow to generate engine thrust, which deteriorates engine performance. The point has been reached in current attempts to increase engine performance by increasing TIT where deterioration of engine performance due to the further increase of turbine cooling air flow overcompensates the cycle benefits the increased TIT would otherwise provide. As such, the desire to achieve future CPRs in the range of 70 for high fan pressure ratio, low bypass ratio engines of the type used for supersonic aircraft and 100 for low fan pressure ratio, high bypass ratio engines of the type used for subsonic aircraft is thwarted.
It is known to use a small amount of the core air prior to delivery to the high pressure compressor of aircraft turbo fan engines for bearing sealing, engine anti-icing, high pressure compressor and low pressure turbine cooling, and the environmental control system ("ECS") or for aircraft cabin air supply. As used herein, this portion of core air flow is referred to as bleed mass flow.